This invention relates to sealing devices and more particularly to a rocket motor joint construction which includes a thermal barrier structure in conjunction with elastomeric primary and secondary O-ring seals between casings of a solid rocket motor.
Assembly joints of current solid rocket motor cases are generally sealed using high performance, elastomeric O-ring seals. The 5500xc2x0 F. propulsion gases commonly produced during the relatively short firing interval of the rocket motors are kept a safe distance away from compounds are used to fill insulation gaps leading to the seals to prevent a flowpath of propulsion gases to the seals.
Normally, these two stages of protection are enough to prevent a direct flowpath of the 900-psi hot gases from reaching the seals. Occasionally, seals have experienced charring due to parasitic leakage paths that open up in the joint-fill compounds during rocket operation. Inspection during disassembly of Space Shuttle solid rocket motor nozzle joints from RSRM-44 and RSRM-45 revealed O-ring erosion of Joint 3 primary O-ring seals. Subsequent improvements in joint-fill compound application techniques have apparently overcome the Joint 3 charring problem. However, a number of nozzle joints including the nozzle-to-case joint and Joint 2 continue to show hot gas penetration through the joint-fill compound. The current nozzle-to-case joint design incorporates primary, secondary, and wiper (innermost) O-rings and polysulfide joint-fill compound. In the current design, one out of seven motors experiences hot gas to the wiper O-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper O-ring results in extensive reviews before resuming flight. Because of the severe conditions which exist in a rocket motor firing environment, further assurance is desired that the joints and the primary and secondary O-ring seals therein are not jeopardized.
It is essential to design rocket motors and their casings to maximize the available thrust from the motor. Solid rocket motors are typically manufactured in sections for assembly in an elongated structure. Not only does the sectional design facilitate manufacture, but the casing sections can be retrieved after firing, refurbished, filled with fuel and reassembled for further firings.
Reusable rocket sections present additional design constraints in that after each firing the casings have experienced considerable stresses and have been exposed to environmental conditions which may change the properties and/or dimensions of the casing structure. The primary and secondary O-ring seals and any additional sealing components must be designed to accommodate such changes. This is in addition to the initial design constraints for the joint structure and the seals therein which may experience relative movement between casing sections due to vibration or the thrust forces produced by the rocket motor.
An additional design constraint imposed upon joint structures including the primary and secondary O-ring seal arrangement is that once the casing sections are assembled, the seals are buried within the structure at a location which is highly tolerant to the physical stresses imposed during motor firing, but not readily accessible to verify the integrity of the seal arrangement. The usual technique for verifying seal integrity is to provide a seal test port which extends between the primary and secondary seals. This test port when pressurized, provides an indication of sealing integrity of the seals under static conditions. However, the condition of materials and structures which make up the mechanical joint between rocket motor casing sections positioned radially inward of the O-ring seals, may have an effect upon pressurization of the seals. It is therefore important that any materials or structures placed radially inward of the O-ring seals must allow gas flow through them so as not to provide misinformation about the integrity of the seals (e.g., false positive) during pressure tests.
Much design and development activity has been directed to finding suitable materials and/or arrangements for the joint configuration between rocket motor casings. Such an arrangement must assure the primary sealing function of the O-ring seals is properly performed and yet accommodate those conditions which are imposed during firing conditions. It has been proposed to include a further thermal barrier structure in the joint between casing sections, radially inward of the primary and secondary O-ring seals. However existing materials have proven unsuitable for this purpose. Thus, there exists a need for a joint and seal structure for casing sections of a rocket motor which provides enhanced protection for the O-ring seals and greater assurance of joint integrity under firing conditions.
It is an object of the present invention to provide an improved thermal barrier, primarily for a rocket motor application.
It is a further object of the present invention to provide a thermal barrier which is used in conjunction with primary and secondary O-ring seals, including wiper seals, in rocket motor casing joints.
It is a further object of the present invention to provide an improved thermal barrier for rocket motors which can withstand rocket motor firing temperatures for a longer period of time.
It is a further object of the present invention to provide an improved thermal barrier for rocket motors that can drop the temperature of incoming jets of hot gas and spread these narrow jets to reduce their damaging effects on downstream O-rings.
It is a further object of the present invention to provide an improved thermal barrier for rocket motors which is of braided configuration and which allows pressurization of the main O-ring seals during rocket motor firing and which does not materially affect main seal integrity pressure tests before rocket firing.
It is a further object of the present invention to provide a casing joint for a rocket motor casing which has enhanced integrity and reliability.
These and other objects of the invention will become apparent in the following Best Modes for Carrying Out Invention and the appended claims.
Rocket motor propulsion gas temperatures reach a level on the order of 5500xc2x0 F. The burn time in typical applications is of the order of only a few minutes. As flows of hot gases occur it is important to prevent such gases from impinging on the O-rings which seal adjacent rocket motor casing sections. Exposure to such hot gases can cause O-ring char and erosion limiting O-ring sealing ability and possibly leading to joint failure, as experienced in the loss of the Space Shuttle Challenger.
Rope seals have been developed previously and braided designs are used in gas turbine engine applications. In these applications the seals are commonly made of ceramic fibers and superalloy wires. These seals provide advantages not only as seals but also as compliant mounts under aggressive temperature and pressure requirements. However, these seals are generally being used in environments at peak temperatures in the 1500-2000xc2x0 F. range. Such seals could not stand up for more than a few seconds to the 5500xc2x0 F. temperatures commonly encountered in rocket motors. This is because ceramics have a melt temperature of about 3500xc2x0 F. and common superalloy metals melt at about 2500xc2x0 F.
It is a teaching of this invention that a braided rope type thermal barrier may be used to achieve advantages in rocket motor type applications. Such a barrier when comprised of carbon fibers provides superior protection to the primary and secondary O-ring seals (including wiper seals) of the motor and additional advantages to the joint structure between motor casing sections. Carbon fibers are used in an exemplary embodiment because of their relatively high heat conduction, low linear expansion coefficient, high corrosion resistance and thermal stability as well as their high strength and low density. It is known that carbon fibers oxidize and lose mass over long intervals when exposed to temperatures above 600-900xc2x0 F. However it has been discovered that carbon fibers formed into a braided rope seal structure are able to withstand very high temperatures for a period of time which is even longer than the burn time of common rocket motors.
A series of tests have been performed to validate the efficacy of the carbon fiber braided rope thermal barrier of the invention including burn tests, temperature drop tests, flow tests, compression tests, and subscale motor tests. The results of the tests indicate that the carbon fiber braided rope thermal barrier of the invention provides adequate resiliency, a very high burn through capability, and a high resistance to hot gas flow. For example, in a 1/5 scale motor test, temperatures over 4200xc2x0 F. were measured on the thermal barrier hot side, while temperatures on the cold side were just under 600xc2x0 F. during rocket firing. This is within the temperature limits of the fluorocarbon rubber (Viton) type material conventionally used for primary and secondary O-rings in rocket motor applications. The thermal barrier of the invention reduces the temperature of incoming hot gas jets, spreads incoming jet flow, and blocks hot slag. The above test results have shown that the thermal barrier provides for great improvement in rocket motor casing joint structure.
The thermal barrier of an exemplary embodiment of the invention comprises essentially an axial carbon fiber central core, surrounded by one or more braid layers of carbon fibers. The thermal barrier is a large, circular, continuous member of relatively small cross sectional dimension, similar to an O-ring. The thermal barrier is disposed in a circular cavity positioned in the joint between rocket motor casing sections, radially inward of the primary and secondary O-ring seals of the motor.
The core of the thermal barrier may comprise a single axial strand or multiple axial strands of carbon fiber material disposed in a generally circular cross-sectional configuration. To facilitate flexibility fewer strands are used in the core. The braid layer or layers radially overlying the core preferably comprise carbon fiber strands surrounding the central core in a braided configuration, each layer sequentially encompassing the central core and any intermediate layers. Preferably, multiple braid layers are employed in lower denier fiber arrangements.
The advantages of such an improved thermal barrier arrangement are many-fold. For example, the use of carbon fiber is highly advantageous in resisting flame or burning at elevated temperatures compared to conventional materials. Thermal barriers of the invention made of carbon material last considerably longer than ceramic, superalloy or elastomeric materials in the severe rocket motor environment.
The structural integrity of the thermal barrier is virtually unchanged in the relatively short time interval in the rocket motor burn cycle during which the thermal barrier may be exposed to large pressure and/or heat transients. This interval is of the order of only several seconds (2-3 seconds) and corresponds to the time for the joint volume between the thermal barrier and the downstream O-rings to fill to the approximate 900 psi pressure within the rocket motor firing chamber. In addition, because of the permeable nature of the carbon thermal barrier structure, pressure equalization occurs across the thermal barrier. Because pressure is equal on both sides of the thermal barrier soon after firing is commenced, the flow of gas through the thermal barrier stops and little, if any, further heat is convected to the O-rings. As a result the O-rings are not subject to charring and erosion, and continue to fulfill their sealing function.
Even after being exposed to the extreme temperatures and pressures of the rocket motor environment, the carbon thermal barrier remains flexible and does not stiffen and/or become brittle as do the prior art ceramic and metal materials.
Formed in a configuration of many braided sheath layers and a relatively small, uniaxial core, the thermal barrier of the preferred form of the invention is more flexible than prior art braided rope configurations. This provides advantages in making it easier to lay and retain the thermal barrier location in the appropriate groove in the joint between casing sections, which facilitates assembly and reduces the chances of damage or misalignment of the thermal barrier. Greater flexibility also makes it easier to ship and store the thermal barrier prior to assembly.
The temperature drop achieved across the thermal barrier of the exemplary embodiment is outstanding. It has been determined in tests that a temperature gradient of the order of 2200xc2x0 F. has occurred across the thermal barrier and has been maintained over relatively long intervals of time, with little or no damage occurring to the barrier structure. The thermal barrier thereby reduces the temperature of gases reaching the primary and secondary O-rings to a level which can be easily accommodated by the O-rings.
Possible heat paths through the thermal barrier have been identified in an effort to account for the large temperature drops that have been observed across the thermal barrier. It is believed that the thermal barrier may act as a xe2x80x9cflame holderxe2x80x9d and trap the thermal mass of the incoming hot gas jet within the braided structure.
As the jets of hot gases pass through the thermal barrier, the heat from these jets can go in several places. Heat from the jet is transferred to the carbon fibers. Some of this heat is conducted along the longitudinal axis of the thermal barrier and away from the heat source. The braided nature of the thermal barrier also spreads out the incoming hot gas jets and causes convection along the length of the thermal barrier. Heat can be transferred to the phenolic walls that surround the thermal barrier in the nozzle joint. This heat can be conducted to the phenolic walls by carbon fibers in the braid, or it can be transferred to the walls by convection from the hot gas that has spread over the thermal barrier. Heat can also be transferred to the phenolic walls through the process of charring the phenolic material. As the phenolic walls become charred, an endothermic reaction, energy is removed from the hot gas jet and the temperature of the gas is reduced. Any small amount of heat remaining in the gas is then transferred out of the thermal barrier and downstream to the O-rings.
Another method in which the thermal barrier acts to drop the temperature across its diameter is through the Joule-Thomson, or throttling, effect. As the hot, high pressure gases pass through small pores between the fibers in the thermal barrier and into the low pressure cavity downstream of the thermal barrier, the gas expands. This expansion causes an additional drop in the temperature of the gas and further reduces the amount of heat that reaches the downstream O-rings. This phenomenon helps to explain how such a large temperature drop occurs across the thermal barrier.
As noted, the improved thermal barrier of the invention has relatively high porosity, which allows leak tests of the primary and secondary O-ring seals to be accomplished without the necessity of a separate seal vent port or other constraints upon the leak test measurement. Further the relatively high porosity feature allows rapid pressurization of the primary and secondary seals under actual motor firing conditions.
While a thermal barrier structure having a generally circular cross-sectional configuration is described as an exemplary embodiment, it will be apparent that other cross-sectional configurations could be employed in other embodiments. Thus, the thermal barrier cross section could as well be square, triangular, rectangular, elliptical or circular, or a combination of these shapes as may be suited for different joint or cavity configurations, environmental conditions or operating conditions such as the need for additional physical support, a longer time interval of protection, a change in porosity or the like.
Various combinations of braid and core configurations can be utilized in embodiments of the invention, as can different numbers of braid layers, to achieve desired properties for the thermal barrier. Plural thermal barrier structures, where the barriers are disposed annularly side by side in the rocket casing joint are also within the teachings of this invention.
In further embodiments of the invention other materials may be added to the thermal barrier combination to achieve various results which enhance certain functions or properties of the thermal barrier. For example, room temperature vulcanized (RTV) sealant can be applied over the outer braid layer or layers of the thermal barrier to enhance the sealing effect and to achieve greater ease of assembly of the rocket motor sections and/or to reduce the porosity characteristics of the thermal barrier.
In still further embodiments of the invention materials such as copper or aluminum can be added to the thermal barrier structure to achieve various desired effects. For example if the core and/or one or more inner braid layers of the thermal barrier is formed of one of such metals, the latent heat of fusion of the metal is utilized so that the metal is sacrificed to maintain a high temperature differential across the thermal barrier. In this way the thermal barrier may withstand prolonged exposure to the hot gases of the rocket motor, providing additional time before the thermal barrier is affected. While molten metal may be deposited within the carbon braids of the thermal barrier, this will also provide the additional advantage of reducing the porosity of the thermal barrier when high porosity of the barrier is no longer desired. That is for example, at a time after initial pressurization of the primary and secondary O-ring seals and when less flow through the thermal barrier is desired, or in circumstances where pressure is being lost through the O-ring seals and it is desirable to minimize flow to prevent erosion and charring of the O-ring seals and possible joint damage. Numerous other advantages of the thermal barrier and joint structure of the invention will be apparent from the following description.